A reduction of the overall drag resistance or total air resistance of an airplane for transportation of goods and/or passengers is in the focus of modern aircraft aerodynamics. For reducing the resistance long lasting investigations have been directed to attempts to keep the airflow in the boundary layer laminar by suitable shapes of the relevant surfaces, also called natural laminar flow or LNF. Such natural laminar flow is provided by a pressure distribution leading to a steady acceleration of the airstream in the region of transonic or supersonic velocities. However, as a side effect, this pressure distribution used for providing natural laminar flow involves a recompression shock wave. The intensity of the recompression shock wave is drastically increased when compared with traditional profiles of flying objects or wings.
This disadvantage of an NLF-profile results both in an increased wave drag and in an increased hazard of separation of the boundary layer at the wing induced by shock waves. A reduction of the shock-wave intensity by influencing the airstream around the flying object is one possible way of avoiding the aforementioned drawbacks.
It is well known that the wave drag increases with an increasing intensity of the shock waves. Furthermore, it is well known that a reduction of the intensity of the shock wave for a given local Mach number of the airstream might be achieved by splitting the shock-wave into some weaker shock waves.
The background art discloses a plurality of different measures and devices for influencing the wave drag at transonic or supersonic wings. Most of the proposed devices are located at the outer surface or integrated into the outer surface of a wing with the aim of splitting the front of a shock wave into a plurality of shock or compression waves and to extend the region of increasing pressure in the airstream near the surfaces. The resulting flow with multiple shock waves has a reduced induced wave drag. One of these known passive measures is to shape a surface in the airstream with local bumps. The bumps cause a softer compression upstream the main shock wave due to the local wall geometry. Other measures integrate cavities or permeable inserts for the airstream into the surface for inducing a passive ventilation of the airstream and for causing a compression upstream of the main shock wave due to the induced secondary flow.
For other active measures a partial flow of the airstream is periodically or continuously sucked through channels of the outer surface of the flying object or air is blown out off these channels into the boundary layer of the airstream. Furthermore, it is well known to actively change the shape of the surface that is positioned in the airstream.
German Patent No. DE 103 05 973 B3 discloses the use of aero-spikes for reducing a wave drag in a transonic airstream, see FIGS. 1 to 3 of this patent. In a supersonic or transonic region bodies generating shock waves are located above a boundary layer near the surface of the flying objects. These bodies might cause an induced separation bubble for causing additional shock waves. The additional shock waves stepwise decelerate the airstream upstream of the main shock wave. The induced shock waves are suitable for influencing a comparatively large region of the airstream in case of the bodies being located with a distance from the surface of the flying object or wing. The increase of pressure in the airstream is not caused abrupt by one single strong shock wave but continuously which leads to a reduction of the overall wave drag.
German Patent No. DE 10 2006 061 709 B3, corresponding to European Patent No. EP 1 939 578 B1, discloses a porous aero-disc. The porous aero-disc is held by an aero-spike at a blunt nose of a supersonic or transonic flying object. The use of a porous aero-disc has been proven to be effective also for variable streaming conditions with improved characteristics when compared with traditional aero-spikes. The aero-disc is built with a material permeable for the airstream comprising a plurality of fine streaming channels for causing a thicker shear layer in a region downstream from the aero-disc.
The present skill for a reduction of a drag resistance and the use of aero-spikes for transonic or supersonic airstreams is summarized in the following articles:                [1] E. Stanewsky et. al.: Drag Reduction by Shock and Boundary Layer Control—Results of the Project EUROSHOCK II, Supported by the European Union 1996-1999, Springer Verlag, 2002        [2] M. Rein, H. Rosemann, E. Schülein: Wave drag reduction by means of aerospikes on transonic wings, In: Hannemann, K.; Seiler, F. (Editors): 26th International Symposium on Shock Waves, CD-ROM Proceedings, ISSW 26, Göttingen (Germany), July 2007.        [3] H. Ogawa, H. Babinsky: Evaluation of wave drag reduction by flow control; Aerospace Science and Technology 10 (2006) 1-8; Elsevier SAS, September 2005.        [4] W. S. Wong, N. Qin, N. Sellars, H. Holden, H. Babinsky: A combined experimental and numerical study of flow structures over three-dimensional shock control bumps; Aerospace Science and Technology 12 (2008) 436-447; Elsevier Masson SAS, October 2007.        [5] B. König, M. Pätzold, T. Lutz, E. Kramer, H. Rosemann, K. Richter, H. Uhlemann: Numerical and Experimental Validation of Three-Dimensional Shock Control Bumps; 4th Flow Control Conference, 23-26 Jun. 2008, Seattle, Wash., AIAA 2008-4001.        